Engine article with ceramic insert and method therefor

ABSTRACT

An article includes a gas turbine engine component that has a ceramic matrix composite (CMC) body and a fiber layer circumscribing a cavity. There is a ceramic insert incorporated into the CMC article body, the insert bordering the cavity.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate ahigh-pressure and temperature exhaust gas flow. The high-pressure andtemperature exhaust gas flow expands through the turbine section todrive the compressor and the fan section. The compressor section mayinclude low and high pressure compressors, and the turbine section mayalso include low and high pressure turbines.

Components in the turbine section are typically formed of a superalloyand may include thermal barrier coatings to extend temperaturecapability and lifetime. Ceramic matrix composite (“CMC”) materials arealso being considered for turbine components. Among other attractiveproperties, CMCs have high temperature resistance. Despite thisattribute, however, there are unique challenges to manufacturing andimplementing CMCs in such components.

SUMMARY

An article according to an example of the present disclosure includes agas turbine engine component that has a ceramic matrix composite (CMC)body that includes a fiber layer circumscribing a cavity, and a ceramicinsert incorporated into the CMC body, the insert bordering the cavity.

In a further embodiment of any of the foregoing embodiments, the gasturbine engine component is an airfoil and the ceramic insert is a ribin the airfoil.

In a further embodiment of any of the foregoing embodiments, the gasturbine engine component is a blade outer air seal and the ceramicinsert is a seal seat in the blade outer air seal.

In a further embodiment of any of the foregoing embodiments, the gasturbine engine component is an airfoil and the ceramic insert is anannular flange in the airfoil.

In a further embodiment of any of the foregoing embodiments, the ceramicinsert is a monolithic ceramic.

In a further embodiment of any of the foregoing embodiments, the fiberlayer includes silicon carbide fibers.

A method according to an example of the present disclosure includesproviding a ceramic insert on a mandrel. The mandrel and the ceramicinsert together define a peripheral working surface, and then forming afiber preform by wrapping a fiber layer around the mandrel and theceramic insert so as to conform to the peripheral working surface. Themandrel is then removed from the fiber preform to leave a cavity in thefiber preform. The ceramic insert remains in the fiber preform andborders the cavity. The fiber preform is then densified with a ceramicmatrix to form a gas turbine engine component.

In a further embodiment of any of the foregoing embodiments, the mandrelis a rod.

In a further embodiment of any of the foregoing embodiments, the mandrelis segmented.

In a further embodiment of any of the foregoing embodiments, the mandrelis multi-piece.

In a further embodiment of any of the foregoing embodiments, the mandrelis of uniform cross-section.

In a further embodiment of any of the foregoing embodiments, thewrapping includes braiding.

In a further embodiment of any of the foregoing embodiments, theremoving of the mandrel includes mechanically removing the mandrel.

A system according to an example of the present disclosure includes amandrel, and a ceramic insert disposed on the mandrel. The mandrel andthe ceramic insert together define a peripheral working surface. A fiberlayer is wrapped around the mandrel and the ceramic insert and conformsto the peripheral working surface.

In a further embodiment of any of the foregoing embodiments, the mandreldefines a central axis, the ceramic insert is located at a first axialposition, and further includes an additional ceramic insert disposed onthe mandrel at a second axial position that is axially-spaced from thefirst axial position.

In a further embodiment of any of the foregoing embodiments, the mandrelis at least one of the following: the mandrel is a rod, the mandrel issegmented, or the mandrel is multi-piece.

In a further embodiment of any of the foregoing embodiments, the fiberlayer is a braided fiber layer.

In a further embodiment of any of the foregoing embodiments, the ceramicinsert is selected from the group consisting of a rib, a seal seat, andan annular flange.

The present disclosure may include any one or more of the individualfeatures disclosed above and/or below alone or in any combinationthereof.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the present disclosure willbecome apparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

FIG. 1 illustrates a gas turbine engine.

FIG. 2 illustrates a portion of a turbine section of the engine.

FIG. 3 depicts an example lay-up process and set-up for fabricating afiber preform of a CMC article.

FIG. 4 illustrates a sectioned view from FIG. 3 .

FIG. 5 depicts another example lay-up process and set-up for fabricatinga fiber preform of a CMC article.

FIG. 6 illustrates a sectioned view from FIG. 5 .

FIG. 7 depicts another example lay-up process and set-up for fabricatinga fiber preform of a CMC article.

FIG. 8 depicts another example lay-up process and set-up for fabricatinga fiber preform for multiple CMC articles.

In this disclosure, like reference numerals designate like elementswhere appropriate and reference numerals with the addition ofone-hundred or multiples thereof designate modified elements that areunderstood to incorporate the same features and benefits of thecorresponding elements.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a housing15 such as a fan case or nacelle, and also drives air along a core flowpath C for compression and communication into the combustor section 26then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in exemplary gas turbine engine20 is illustrated as a geared architecture 48 to drive a fan 42 at alower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high) pressure turbine 54. Acombustor 56 is arranged in the exemplary gas turbine 20 between thehigh pressure compressor 52 and the high pressure turbine 54. Amid-turbine frame 57 of the engine static structure 36 may be arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 57 further supports bearing systems 38in the turbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded through the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aftof the combustor section 26 or even aft of turbine section 28, and fan42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), andcan be less than or equal to about 18.0, or more narrowly can be lessthan or equal to 16.0. The geared architecture 48 is an epicyclic geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3. The gear reduction ratio maybe less than or equal to 4.0. The low pressure turbine 46 has a pressureratio that is greater than about five. The low pressure turbine pressureratio can be less than or equal to 13.0, or more narrowly less than orequal to 12.0. In one disclosed embodiment, the engine 20 bypass ratiois greater than about ten (10:1), the fan diameter is significantlylarger than that of the low pressure compressor 44, and the low pressureturbine 46 has a pressure ratio that is greater than about five 5:1. Lowpressure turbine 46 pressure ratio is pressure measured prior to aninlet of low pressure turbine 46 as related to the pressure at theoutlet of the low pressure turbine 46 prior to an exhaust nozzle. Thegeared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.3:1 and less than about 5:1. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption ('TSFC)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. The engine parameters described above and those in thisparagraph are measured at this condition unless otherwise specified.“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45, or more narrowly greater than orequal to 1.25. “Low corrected fan tip speed” is the actual fan tip speedin ft/sec divided by an industry standard temperature correction of[(Tram ° R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed” asdisclosed herein according to one non-limiting embodiment is less thanabout 1150.0 ft/second (350.5 meters/second), and can be greater than orequal to 1000.0 ft/second (304.8 meters/second).

FIG. 2 illustrates a portion of the turbine section 28 of the engine 20.The turbine section 28 includes a row of vanes 60 and a row of blades62. The row of vanes 60 and the adjacent row of blades 62 togetherconstitute a stage. The turbine section 28 may include multiple stages.Each vane 60 includes inner and outer platforms 64/66 and an airfoilsection 67 that extends between the platforms 64/66. The vane 60 ishollow and has a cavity 68 that extends there through for conveyingcooling air to downstream structures, such as tangential onboardinjectors (TOBIs). Each blade 62 includes a platform 70 and an airfoilsection 72 that extends there from. The airfoil section 72 has aninternal cavity 74.

There is a row of blade outer air seals (BOAS) 76 arranged near the tipsof the blades 62. Each BOAS 76 includes a cavity 78 for conveyingcooling air. The BOAS 76 may be supported from an engine turbine case orother static structure. The BOAS 76 provides sealing at the tips of theblades 62 to reduce the leakage flow of combustion gases around the tipsof the blades 62. The constituent portions of each of the vane 60, blade62, and BOAS 76 above constitute the body of each of these articles. A“cavity” as used herein may include a relatively large volume, used forbulk cooling air flow for example, or a relatively smaller passage orpassages, used for cooling portions of the airfoil section 72 forexample. It is to be appreciated that the vane 60, blade 62, and BOAS 76are shown schematically and may include additional structures that areknown to those of ordinary skill in the art.

The body of any or each of the vane 60, the blade 62, and the BOAS 76 isformed from a ceramic matrix composite (CMC) material. It is also to beappreciated that although the examples herein are presented in contextof the vane 60, the blade 62, and/or the BOAS 76, the examples are alsoapplicable to other gas turbine engine components that are formed ofCMC. A CMC material is comprised of a ceramic reinforcement, which isusually continuous ceramic fibers, in a ceramic matrix. Example ceramicmatrices are silicon-containing ceramic, such as but not limited to, asilicon carbide (SiC) matrix or a silicon nitride (Si3N4) matrix.Example ceramic reinforcements of the CMC are silicon-containing ceramicfibers, such as but not limited to, silicon carbide (SiC) fiber orsilicon nitride (Si3N4) fibers. The CMC material may be, but is notlimited to, a SiC/SiC ceramic matrix composite in which SiC fiber layersare disposed within a SiC matrix. The fiber layers have a fiberarchitecture, which refers to an ordered arrangement of the fiber towsrelative to one another, such as a 2D woven layer or a 3D structuredlayer. Unlike a CMC, a monolithic ceramic does not contain fibers orreinforcement and is formed of a single material, such as monolithicsilicon carbide (SiC) or monolithic silicon nitride (Si3N4).

Hollow CMC gas turbine engine components may be fabricated in a lay-upprocess that involves arranging one or more fiber layers around amandrel that replicates the geometry of the cavity in the component. Forexample, first a fiber preform is formed by braiding fiber layers aroundthe mandrel. The mandrel is then removed, leaving the cavity and thebraided fiber layers encompassing the cavity. The preform is thensubjected to additional processes to produce the end-use article. Forinstance, portions of the preform may be re-shaped to a desired designgeometry and/or to accommodate installation of inserts, the preform maybe trimmed of excess fibers, or sub-components may be installed into thepreform.

Each additional process that the preform is subjected to involveshandling the preform in some manner, which increases the chances thatthe geometry of the preform from the braiding step will be inadvertentlyaltered and deviate from the desired design geometry. Therefore, to theextent that an additional process can be eliminated, the chances ofaltering the geometry will be reduced. Nevertheless, this is easier saidthan done, as it is challenging to eliminate steps while still meetingperformance and manufacturability requirements imposed by the articledesign. Along these lines, as disclosed herein, the elimination of oneor more additional processing steps for such fiber preforms may befacilitated by incorporating a ceramic insert into the preform at thefiber layer lay-up stage.

FIG. 3 illustrates representative features of a fiber layer lay-upprocess for fabricating an airfoil, such as the vane 60 or the blade 62.FIG. 4 is a sectioned view of FIG. 3 . The process involves a mandrel 80that may be formed of graphite, but is not limited thereto. In thiscase, the mandrel 80 is a multi-piece mandrel that includes a firstmandrel rod piece 80 a and a second mandrel rod piece 80 b. Thecross-sectional geometries of the mandrel rod pieces 80 a/80 b may bethe same or different from each other and correspond to the desireddesign geometry of the cavity 68/74 that is to be formed. In thisregard, the cross-sections may be, but are not limited to, oval,circular, stadium shape, wing shape, portion of a wing shape, or otherdesired geometry of the cavity 68/74 or sub-cavities that make up thecavity 68/74. The cross-sections of each mandrel rod piece 80 a/80 b maybe constant from end-to-end or tapered from one end or intermediatelocation to the other end or other intermediate location.

The cavity 68/74 is to be divided into sub-cavities by a rib thatextends from one side of the airfoil to the other, e.g., from a pressureside to a suction side. Ribs may sometimes be formed after a lay-upprocess by adding structures to the fiber preform prior to densificationwith the matrix material, or by adding filler material (oftencolloquially referred to as a noodle) along the lengths of the mandrelrod pieces. As indicated above, processing after formation of the fiberpreform increases the chances of inadvertently altering the geometry.Moreover, filler materials may shift during handling or processing andultimately cause the rib to deviate from its design position and/ordesign geometry. In this regard, rather than a filler material or anadditional process after lay-up, a ceramic insert 82 is used between themandrel rod pieces 80 a/80 b.

The ceramic insert 82 may be a pre-fabricated from a CMC or a monolithicceramic. The cross-sectional geometry of the ceramic insert 82 in thisexample has a central wall 82 a and flared ends 82 b/82 c (FIG. 4 ). Theflared ends 82 b/82 c are substantially flush against the mandrel rodpieces 80 a/80 b. Optionally, the mandrel rod pieces 80 a/80 b andceramic insert 82 may be held together using a fixture or adhesivematerial, such as a tackifier, wax, or preceramic polymer. Tackifiersand waxes may later burn off during thermal cycles used fordensification, while preceramic polymer thermally converts to ceramic.

The periphery formed by the outline of the mandrel rod pieces 80 a/80 band the ceramic insert 82 defines a working surface 83 around which oneor more fiber layers 84 will be wrapped to form a fiber preform 86. Forexample, the fiber layer or layers 84 are braided (e.g., triaxialbraiding) around the mandrel rod pieces 80 a/80 b and ceramic insert 82and conform to the working surface 83 so as to be in close intimatecontact with the periphery surfaces of the mandrel rod pieces 80 a/80 band the ceramic insert 82. As will be appreciated, the fiberarchitecture of the fiber layer or layers 84 herein are not limited tobraiding. As used herein, the arrangement of a mandrel, a fiber layer(s)on the mandrel, and a ceramic insert on the mandrel is considered to bea “system,” i.e., a system representing a set-up for fabrication of CMCarticles, such as the vane 60, the blade 62, or the BOAS 76.

The mandrel rod pieces 80 a/80 b are then removed from the fiber preform86 after braiding. The technique by which the mandrel rod pieces 80 a/80b are removed is not particularly limited. Most typically, the mandrelrod pieces 80 a/80 b will be removed mechanically by pulling the pieces80 a/80 b out from the fiber preform 86. Other techniques, such as butnot limited to, chemical removal and thermal removal may alternativelybe used or be used in combination. The ceramic insert 82 remains in thefiber preform 86 after removal of the mandrel rod pieces 80 a/80 b andborders the cavity left by the mandrel rod pieces 80 a/80 b.

The fiber preform 86 may be subjected to one or more additionalprocesses prior to densification, such as trimming steps to removeexcess fibers or forming operations to shape portions of the fiberpreform (e.g., to form one or more platforms, flanges, etc.). However,no further process are required to install a rib because the ceramicinsert 82 remains in the fiber preform 86 and serves as the rib. Thefiber preform 86 (with the ceramic insert 82) is then subjected todensification in which the ceramic matrix provided. Although notlimited, the densification process may include chemical vaporinfiltration, polymer infiltration and pyrolysis or combinations ofthese and/or other ceramic processes.

In addition to facilitating elimination of steps for forming the ribafter formation of the fiber preform 86, the ceramic insert 82 may alsoprovide surfaces that act as seal seats for sealing. For instance, theceramic insert 82 can be prefabricated with a high density, low surfaceroughness, and/or precision geometry to enhance sealing. A high densityreduces surface porosity and thus facilitates the provision of a smoothsurface to seal against. The ceramic insert 82 may also be polished orcoated to form smooth surfaces to seal against. Moreover, the ceramicinsert 82 may be fabricated to high dimensional tolerances to enhancerelative positioning of the seal seat. If a ceramic insert were insteadinstalled into the fiber preform after forming the fiber preform, theinstallation may increase the chances of inadvertently altering thegeometry of the fiber preform. Likewise, machining processes afterformation of the fiber preform that are used to create or smooth outseal seats may also increase the chances of inadvertently altering thegeometry of the fiber preform. The use of the ceramic insert 82 thusfacilitates elimination of those processes.

FIG. 5 illustrates representative features of another example fiberlayer lay-up process for fabricating an airfoil, such as the vane 60 orthe blade 62. FIG. 6 is a sectioned view of FIG. 5 . The processinvolves a mandrel 180 (rod) and the ceramic insert 182 is an annularflange. The mandrel 180 extends through the open middle of the annularflange, which may be held on the mandrel 180 as discussed above. Thefiber layer or layers 84 are braided around the mandrel 180 and at leasta portion of the ceramic insert 182 and conform to the working surface83 so as to be in close intimate contact with the periphery surfaces ofthe mandrel 180 and the ceramic insert 182.

The mandrel 180 is then removed from the fiber preform 86, as discussedabove. The ceramic insert 182 remains in the fiber preform 86 afterremoval and borders the cavity left by the mandrel 180. The fiberpreform 86 is densified, as also discussed above. The ceramic insert 182facilitates elimination of steps for forming the annular flange afterformation of the fiber preform 86. The annular flange may also providesurfaces that act as seal seats for sealing. For instance, the annularflange is located at a radial end of the airfoil, e.g., an inlet to thecavity in the airfoil and one or more seals may be provided in contactalong the surface of the annular flange. An additional annular flange(additional ceramic insert 182) may also be used at the other radialend, e.g., outlet of the cavity in the airfoil. In further examples, theannular flange may also mate with a sub-component that is installed inthe airfoil. As an example, a tube may be provided through the cavity ofthe airfoil, and the annular flange may contact and/or seal with thetube. The enhanced sealing and/or mating with the tube facilitateslowering less pressure loss of cooling air flow and, as a result,efficiency improvements.

The ceramic insert 182 may thus be prefabricated with a high density,low surface roughness, and/or precision geometry to enhance sealingand/or mating with a sub-component. If a ceramic insert were insteadinstalled into the fiber preform after forming the fiber preform, theinstallation may increase the chances of inadvertently altering thegeometry of the fiber preform. Likewise, machining processes afterformation of the fiber preform that are used to create or smooth outseal seats or contact surfaces may also increase the chances ofinadvertently altering the geometry of the fiber preform.

FIG. 7 illustrates representative features of another example fiberlayer lay-up process for fabricating the BOAS 76. The process involves amandrel 280 that is segmented into first and second mandrel segments 280a/280 b. The segments 280 a/280 b differ from the multi-pieces 80 a/80 bdescribed above for the mandrel 80 in that the segments 280 a/280 b areof complementary geometries that fit together to form a single mandrel,while the multi-pieces 80 a/80 b are of non-complementary geometriesthat do not fit together. In the example shown, the axial ends of thesegments 280 a/280 b fit together to form a single, elongated mandrel.

One or more ceramic inserts 282 are disposed on the mandrel 280 and maybe held in place as discussed above. The fiber layer or layers 84 arebraided around the mandrel 280 and the ceramic inserts 282 and conformto the working surface 83 so as to be in close intimate contact with theperiphery surfaces of the mandrel 280 and the ceramic insert 282.

The mandrel 280 is then removed from the fiber preform 86. In this case,the segments 280 a/280 b are pulled in opposite directions so as to beremoved from the ends of the fiber preform 86. The ceramic inserts 282remain in the fiber preform 86 after removal and border the cavity leftby the mandrel 280. The fiber preform 86 is then densified, as discussedabove. The ceramic inserts 282 may serve as seal seats and/or asattachment support features for installation of the BOAS 76 into theengine 20. The ceramic inserts 282 facilitate elimination of steps forforming seal seats or attachment features after formation of the fiberpreform 86. The ceramic inserts 282 may thus be prefabricated with ahigh density, low surface roughness, and/or precision geometry toenhance sealing and/or geometry for contact with mating sub-components.If a ceramic insert were instead installed into the fiber preform afterforming the fiber preform, the installation may increase the chances ofinadvertently altering the geometry of the fiber preform. Likewise,machining processes after formation of the fiber preform that are usedto create or smooth out seal seats or attachment features may alsoincrease the chances of inadvertently altering the geometry of the fiberpreform.

Use of a segmented mandrel also enables a plurality of BOAS 76 articlesto be fabricated simultaneously, in one wrapping step. For instance, inFIG. 8 the mandrel 380 is segmented into a plurality of mandrel segments380 a/380 b/380 c (three shown, but additional segments may be used). Inthis case, each of the segments 380 a/380 b/380 c is of uniformcross-section along its entire length. The segments 380 a/380 b/380 care arranged end-to-end. Two or more ceramic inserts 382 are disposedalong the mandrel 380. In particular, one of the ceramic inserts 382 islocated at a first axial position L1 and the other ceramic insert 382 islocated at a second axial position L2 that is axially-spaced from thefirst axial position L1. The positions L1 and L2 correspond to thelocations of the ends of the fiber preforms 86 a/86 b/86 c that will beproduced. The fiber layer or layers 84 are then braided around themandrel 380 and the ceramic inserts 382 and conform to the workingsurface 83 so as to be in close intimate contact with the peripherysurfaces of the mandrel 380 and the ceramic inserts 382.

The mandrel 380 is then removed from the fiber preforms 86 a/86 b/86 c,which at this stage are connected as a braided uni-sleeve. For instance,the segments 380 a/380 b/380 c are pulled in axial directions so as tobe removed from the ends of the uni-sleeve. The ceramic inserts 382remain in the uni-sleeve after removal and border the cavity left by thesegments 380 a/380 b/380 c. The uni-sleeve may then be densified, asdiscussed above. After densification, the densified uni-sleeve is cut atthe locations L1/L2 to form three BOAS 76 articles. Alternatively, theuni-sleeve may be subjected to a low temperature consolidation step toform a brown uni-sleeve, which is then cut at the locations L1/L2 toform three brown BOAS 76 articles. The brown BOAS article are thenprocessed to full densification. In this manner, multiple BOAS 76articles can be processed in a single wrapping step, rather than threeseparate wrapping steps to form each individual fiber preform.

Although a combination of features is shown in the illustrated examples,not all of them need to be combined to realize the benefits of variousembodiments of this disclosure. In other words, a system designedaccording to an embodiment of this disclosure will not necessarilyinclude all of the features shown in any one of the Figures or all ofthe portions schematically shown in the Figures. Moreover, selectedfeatures of one example embodiment may be combined with selectedfeatures of other example embodiments.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthis disclosure. The scope of legal protection given to this disclosurecan only be determined by studying the following claims.

1. An article comprising: a gas turbine engine component having aceramic matrix composite (CMC) body including a fiber layercircumscribing a cavity, and a ceramic insert incorporated into the CMCbody, the insert bordering the cavity wherein the ceramic insert is anannular flange, and the fiber layer is braided around the annular flangeso as to be in contact with, and conform to, periphery surfaces of theannular flange. 2-4. (canceled)
 5. The article as recited in claim 1,wherein the ceramic insert is a monolithic ceramic.
 6. The article asrecited in claim 5, wherein the fiber layer includes silicon carbidefibers.
 7. A method comprising: providing a ceramic insert on a mandrel,the mandrel and the ceramic insert together define a peripheral workingsurface, the ceramic insert is an annular flange; forming a fiberpreform by wrapping a fiber layer around the mandrel and the ceramicinsert so as to conform to the peripheral working surface; removing themandrel from the fiber preform to leave a cavity in the fiber preform,the ceramic insert remaining in the fiber preform and bordering thecavity; densifying the fiber preform with a ceramic matrix to form a gasturbine engine component.
 8. The method as recited in claim 7, whereinthe mandrel is a rod.
 9. The method as recited in claim 7, wherein themandrel is segmented.
 10. The method as recited in claim 7, wherein themandrel is multi-piece.
 11. The method as recited in claim 7, whereinthe mandrel is of uniform cross-section.
 12. The method as recited inclaim 7, wherein the wrapping includes braiding.
 13. The method asrecited in claim 7, wherein the removing of the mandrel includesmechanically removing the mandrel. 14-18. (canceled)
 19. The article asrecited in claim 1, wherein the annular flange includes a lip thatprotrudes from the CMC body.
 20. The article as recited in claim 1,wherein the gas turbine engine component is an airfoil.
 21. An articlecomprising: a gas turbine engine component having a ceramic matrixcomposite (CMC) body including a fiber layer circumscribing a cavity,and a ceramic insert incorporated into the CMC body, the insertbordering the cavity.